This invention relates to the field of solid fuel rocket motors and the controlling of thrust in solid fuel rocket motors through the use of electrically controlled fuel preheating.
Solid fuel rockets are preferred for military and long mission uses because of their inherent simplicity and their avoidance of the complex plumbing, mixing and control elements required in liquid fuel rockets and because of the ease and safety with which the rocket fuel, or grain, can be preserved for future use. The design parameters, fuel selection tradeoffs and operating characteristics of solid fuel rockets are well known in the art and are, for example, discussed in the reference textbooks “Rocket Propulsion Elements”, 2d ed., by G. P. Sutton, and “Propellant Chemistry” by Stanley F. Sarner.
A characteristics of solid fuel rockets is the difficulty involved in modulation or termination of the thrust-producing reaction once grain burn has commenced. The need for thrust modulation or thrust termination plus re-initiation is readily apparent in practical applications of a rocket in military or scientific uses. The functions of threat avoidance, multiple purpose missions and vehicle atmospheric reentry each present a desirable environment for some form of thrust change, for example.
Questions of complete burn termination and re-ignition are not considered in this specification; present concerns are with thrust modulation or control during the burning of solid fuel rocket grain. Such thrust modulation might, for example, also be desirable in tailoring the placement velocity and the orbit of a spacecraft, in trading thrust magnitude for thrust duration in a rocket, or in balancing the thrust applied to a multiple rocket vehicle, especially during the initial liftoff, low air velocity, flight portion.
The achievement of practical and workable thrust modulation in a solid fuel rocket has been attended by a significant degree of difficulty, however, and in most heretofore practiced arrangements has required significant compromises with respect to rocket payload, reliability, attained degree of control, and added rocket complexity. The variable exit nozzle geometry arrangements previously used for thrust control are, for example, found to be large in physical size, heavy and weight penalizing with respect to rocket payload, requiring of specialized high-temperature, high-strength materials, and productive of large pressure excursions within the propellant combustion chamber. Such nozzle geometry changing arrangements or throttles are also found to be of limited compatibility with the gimbaled or pivoting nozzle arrangements that are commonly employed for guidance and stability control of a rocket and its payload. The pulse modulated solid fuel rocket propellant modulators, although somewhat successful over a limited burn rate schedule, are limited in control ability over the desired larger range of thrust changes and additionally have the characteristic of ejecting unburned inhibitor media. The moving filament throttle arrangements are also cumbered by large weight and space requirements for the attending winch mechanism and are attended by unresolved problems of binding and fusing throttle filament elements so as to be for all practical purposes, not presently desirable.
The use of burn rate increasing energy communicated to local areas of the rocket grain adjacent the burn face in order to produce coning action and increased burn face surface area, is suggested in the patents of Winch et al, U.S. Pat. No. 4,587,805, and DeHaye, U.S. Pat. No. 3,529,425 and presently affords one of the more practical arrangements for controlling solid fuel rocket grain burn rates. The present invention adds a new dimension to this class of burn rate control by additionally simplifying the required controlling apparatus and limiting the mass and volume of control apparatus attending the rocket structure.